A gas turbine engine generally includes a fan and a core arranged in flow communication with one another. Additionally, the core of the gas turbine engine generally includes, in serial flow order, a compressor section, a combustion section, a turbine section, and an exhaust section. In operation, air is provided from the fan to an inlet of the compressor section where one or more axial compressors progressively compress the air until it reaches the combustion section. Fuel is mixed with the compressed air and burned within the combustion section to provide combustion gases. The combustion gases are routed from the combustion section to the turbine section. The flow of combustion gases through the turbine section drives the turbine section and is then routed through the exhaust section, e.g., to atmosphere.
A typical stage of rotor blades, such as a turbine rotor blade stage, includes a shroud positioned radially outward from a platform of each blade, near the tips of the blades. Similarly, a typical stage of stator vanes or nozzles, such as a turbine nozzle stage, includes an outer band positioned radially outward from an inner band, at the radially outer end of each nozzle. Accordingly, at the axial interface between adjacent blade and nozzle stages, a gap exists between the shroud of the blade stage and the outer band of the nozzle stage. As such, fluid flowing within or around the stages, such as combustion gases flowing through the stages of the turbine section, may leak through the gap between the shroud and the outer band of the nozzle stage, which can impact engine performance. The axial shroud-outer band interface may pose other problems as well.
Therefore, an improved interface between a rotor blade stage and a nozzle stage of a gas turbine engine would be desirable. In particular, a shroud that extends over both a rotor blade stage and a nozzle stage, e.g., such that the shroud forms the outer band of the nozzle stage, would be advantageous. More particularly, a shroud over a rotor blade stage that extends axially aft through an adjacent nozzle stage with openings for nozzles of the nozzle stage would be beneficial. In addition, features for sealing the nozzles inserted through the openings in the shroud to seal the nozzle and the shroud would be desirable. Moreover, a shroud and/or nozzles formed from a ceramic matrix composite (CMC) material would be useful. Further, a method of assembling a gas turbine engine to include a shroud that forms an outer wall of both a rotor blade stage and a nozzle stage would be desirable.